Gas turbine engines (“GTE”) have been engineered extensively to improve efficiency, thrust-to-weight ratios, and other measures of engine performance. One of the most direct manners by which engine performance can be improved is through increases in the core rotational speeds and turbine inlet temperatures generated during engine operation. However, as turbine inlet temperatures and rotational speeds increase, so too do the thermal and mechanical demands placed on the GTE components. The most demanding performance requirements are typically placed on the high-pressure turbine rotor or rotors, which are positioned immediately downstream of the GTE combustion section and which rotate at the greatest speed during engine operation. The turbine blades, as well as surface regions of the turbine disk, in particular, are directly exposed to combustive gas flow at or near peak temperatures and are consequently heated to exceedingly high temperatures at which most alloys weaken or melt and become prone to oxidation or other forms of chemical and structural degradation.
Turbines can be broadly divided into two categories, axial and radial turbines, based upon the direction of airflow received by the turbine relative to the turbine's rotational axis. Each type of turbine has benefits and tradeoffs. For example, relative to axial turbines, radial turbines offer certain performance benefits including reduced aerodynamic loading which enable the turbine to operate at greater efficiencies, and higher tip speeds which reduce relative total temperatures that allow the turbine to operate at higher temperatures. However, due to the nature of their design, radial turbines are relatively lengthy in the axial direction. As a result, some radial turbines can be undesirably heavy and difficult to cool at the disk rim, especially at the mid section. Axial turbines are lighter weight due to shorter axial length but has relatively higher aerodynamic loading. So despite having shorter conduction distance from the sides of the disk hub to the rim, the higher gas path velocities can overwhelm the conduction cooling capability of some axial turbines. Regardless of the configuration, however, present turbine cooling schemes are unable to adequately cool the surface of the disk (between the blades, commonly referred to in the art as the “throat” region) of the turbine, resulting in undesirable high component metal operating temperatures and temperature gradients, especially during start-up and transient operational conditions. For at least these reasons, and despite the proposal of multiple axial and radial turbine designs in the prior art, few currently-implemented gas turbine engine platforms are able to operate at optimally-high temperatures without risking melting, oxidation, and/or other forms of degradation.
It would thus be desirable to provide a turbine suitable for usage in a gas turbine engine that can operate at elevated turbine inlet temperature levels. It would further be desirable to provide a turbine that has improved inter-blade disk surface region cooling characteristics. Furthermore, other desirable features and characteristics of the invention will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.